This invention relates generally to modular spacecraft engineering and, more particularly, to spacecraft electronic equipment modules that can be externally mounted on a core spacecraft structure. Equipment panels in conventional spacecraft have had high vibration responses due to acoustic conditions during launch, launch transients transmitted through the spacecraft structure, or both. These high vibration responses have adversely affected electrical and mechanical components mounted on the panels, and the panel structure itself. No serious attempt has been made to minimize the vibration response. Instead, the electrical components were designed to be compatible with whatever vibration environment the panels provided. It will be helpful to review, by way of further background, the nature of electronic equipment panel design for spacecraft prior to the present invention.
Conventional spacecraft designs have placed electronic equipment on panels located inside box-like modules, with the panels typically forming the walls of the modules. The modules are mounted onto a spacecraft and interconnected with complex interconnecting wiring. Heat dissipates from the electronics equipment into the mounting and radiates into space. The size of each module is determined in part by the heat radiating area needed to cool the equipment, so there is often unused volume within each module box or compartment. Because only one side of the equipment panel is used as a radiator, the modules often have to be very large to maintain desired operating temperatures. A related problem is that accessing equipment requires the removal of panels from the equipment compartment. Not only are modules of this type bulky to accommodate in a launch vehicle and difficult to access for servicing, but they typically need to be coupled to other thermal radiator panels, through heat-conducting pipes, to provide overall thermal management of the spacecraft.
Further, because the box modules are rigidly bolted to the core structure of the spacecraft, thermally induced stresses are a significant problem because of temperature differences between the modules and the spacecraft core structure. In brief, these conventional equipment modules are structurally and thermally dependent on the spacecraft core structure, and the overall design of the spacecraft must take into account the thermal requirements of each module and the structural forces resulting from the presence of each module.
Because of the design limitations of prior equipment panels for spacecraft, and the manner in which they have been housed and mounted, no serious consideration has been given to managing the vibration environment of the equipment. The vibration environment has been aggravated by the use of rigid, statically indeterminate mounting of the equipment to the spacecraft core structure. Launch transients are inevitably transmitted to the electronic equipment through this type of mounting. Further, the use of relatively large equipment panels for appropriate thermal management has resulted in high vibration responses to acoustic energy generated at launch. It will be appreciated from the foregoing that there is a need for improvement in the area of vibration management in spacecraft, especially as it concerns electronic equipment panels. The present inventions addresses this problem and has additional advantages that will become apparent from the following summary.